Historically, satellite architecture has featured a small number of relatively large spacecraft carrying all instrumentation required to carry out prescribed missions. There is a current trend in some applications, however, toward deployment of a multiple satellite, or “multi-sat” approach, such as envisioned in NASA's GOES-R system architecture featuring smaller spacecraft each incorporating different subsets of instrumentation. The “multi-sat” architecture is advantageous because of its increased implementation and operational flexibility since instruments may be launched as they become available; there is no need to delay launches and system implementation until all instruments are available. Greater flexibility also exists by this architecture with regard to in-orbit sparing, because orbital locations of primary instruments can adjusted independently to minimize the coverage impacts of spacecraft or instrument failures. Furthermore, multi-sat architecture lends itself to a spiral upgrade approach using a common scalable bus design with substantial growth margin. The design scalability eliminates the need for expensive spacecraft redesigns or block changes as system requirements change and new instruments are introduced.
A drawback of multi-sat architecture is that it requires at least twice as many spacecraft to implement, compared to current “large-sat” architecture. If the spacecraft are so heavy that they require dedicated launches, then the increased launch cost might render the multi-sat architecture too expensive to implement. Therefore, to provide a cost-effective system solution, the spacecraft must be dual-launch compatible.
To enable a dual launch with payload mass margin drives spacecraft designers to consider a standard dual-wing solar array configuration. Such a configuration, which is typical for GEO communications spacecraft, has one solar array that extends from the north-facing panel and another that extends from the south-facing panel. This “balanced” configuration tends to result in small solar pressure disturbance torque and minimal propellant and operational down time needed for momentum control thruster maneuvers. With less propellant needed for momentum adjust, greater payload mass is accommodated while maintaining the desired dual launch compatibility.
An additional benefit of a dual-wing configuration is in reduced reaction wheel (RWA) operating speeds and jitter, and reduced solar array thermal snap attitude disturbance at eclipse exit. A disadvantage, however, is that one of the solar arrays will impede the field of-view (FOV) of the instrument passive cryogenic radiators, which prevents achieving the required temperature control of critical focal plane sensor elements. If instrument focal plane temperatures cannot be maintained, then mission performance will be impacted. Therefore, despite the advantage of a dual-array configuration, current spacecraft tend to feature a single solar array wing. The single wing design is used because it leaves one side of the spacecraft (either the north or south side) with an unobstructed view of cold space. The instruments are then mounted on the spacecraft's earth facing panel such that their thermal radiators take advantage of this unobstructed view, and hence their focal planes may be maintained at the proper temperatures. The single wing configuration tends to lead to a large solar pressure torque, large momentum-adjust propellant requirements, frequent (e.g., daily) momentum adjust maneuvers that disrupt instrument operations, and higher RWA speeds that produce larger disturbances and jitter. And the single wing design considerably limits dual launch payload mass capability because of the need to carry a large amount of momentum adjust propellant.
Another disadvantage of current GOES architecture is that each instrument must include its own thermal radiators and active coolers as necessary to maintain the proper focal plane temperatures. For passively cooled instruments with cryogenic thermal requirements the required radiator size may be large (e.g., more than 10 ft2) and the need to accommodate such radiators may drive the instruments to be larger and heavier than they would nominally have to be. Also, a GEO spacecraft design includes its own radiators (e.g. the north and south-facing panels), and these radiators are generally underutilized for a GEO remote sensing mission due to overall low thermal dissipation requirements. The fact that the instruments do not have access to these radiators, and hence must supply their own duplicate radiators, is highly mass inefficient.
Instrument radiators are located on only one side of the instrument, which either faces south or north. To prevent instrument radiator sun exposure, GEO remote sensing spacecraft must execute a 180-degree yaw flip every 6 months. During the yaw flip and the subsequent settling transient, the instrument data is unavailable.
Additionally, the present approach to instrument thermal control lacks scalability. If additional instrument heat must be rejected, then the instruments must grow in size and mass to accommodate the larger radiators or active coolers. Also, the temperature of the focal planes of passively cooled instruments may fluctuate due to daily and seasonal changes in the instrument thermal environment. Such temperature variations may possibly degrade measurement accuracy.
Finally, next generation primary instruments (i,e., imagers and sounders) have extremely stringent pointing and jitter requirements. Active coolers incorporated in the instruments may produce mechanical vibrations that can impact the measurement accuracy. Mitigation of these vibration effects may complicate instrument designs and increase cost.